It is known to use ceramic matrix composite (CMC) materials in high temperature environments such as in the hot combustion gas flow path of a gas turbine engine. CMC materials offer the potential for higher operating temperatures than do metal alloy materials due to the inherent nature of ceramic materials. This capability may be translated into a reduced cooling requirement that, in turn, may result in higher power, greater efficiency, and/or reduced emissions from the machine. However, CMC materials generally are not as strong as metal, and therefore the required cross-section for a particular application may be relatively thick. Due to the low coefficient of thermal conductivity of CMC materials and the relatively thick cross-section necessary for many applications, backside closed-loop cooling is generally ineffective as a cooling technique for protecting these materials in combustion turbine applications. Accordingly, high temperature insulation for ceramic matrix composites has been described in U.S. Pat. No. 6,197,424 B1, which is commonly assigned with the present invention and is incorporated by reference herein. That patent describes an oxide-based insulation system for a ceramic matrix composite substrate that is dimensionally and chemically stable at a temperature of approximately 1600° C. That patent also describes a stationary vane for a gas turbine engine formed from such an insulated CMC material. A similar gas turbine vane 10 is illustrated in FIG. 1 herein as including an inner wall 12 and stiffening ribs 14 formed of CMC material covered by an overlying layer of insulation 16. Backside cooling of the inner wall 12 is achieved by convection cooling, e.g. via direct impingement through supply baffles (not shown) situated in the interior chambers 18 using air directed from the compressor section of the engine.
If baffles or other means are used to direct a flow of cooling fluid throughout the airfoil member for backside cooling and/or film cooling, the cooling fluid is typically maintained at a pressure that is in excess of the pressure of the combustion gasses on the outside of the airfoil so that any failure of the pressure boundary will not result in the leakage of the hot combustion gas into the vane. Such cooling passages must generally have a complex geometry in order to provide a precise amount of cooling in particular locations to ensure an adequate degree of cooling without over-cooling of the component. It is generally very difficult to form such complex cooling passages in a ceramic matrix composite component. Alternatively, large central chambers 18 as illustrated in FIG. 1 may be used with appropriate baffling to create impingement of the cooling fluid onto the backside of the surface to be cooled. Such large chambers create an internal pressure force that can result in the undesirable ballooning of the airfoil structure due to the internal pressure of the cooling fluid applied to the large internal surface area of the passage 18. Furthermore, the geometry of FIG. 1 is also limited by stress concentrations at the intersection of the stiffening ribs 14 and the innerwall 12.